Electric propulsion apparatus

ABSTRACT

An electric propulsion machine includes an ion thruster having an annular discharge chamber housing an anode having a large surface area. The ion thruster includes flat annular ion optics with a small span to gap ratio. Optionally, a second electric propulsion thruster may be disposed in a cylindrical space disposed within an interior of the annulus.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority to U.S. Provisional Application Ser.No. 61/295,326 filed Jan. 15, 2010.

ORIGIN OF THE INVENTION

The invention described herein was made by employees of the UnitedStates Government and may be manufactured and used by or for theGovernment for Government purposes without the payment of any royaltiesthereon or therefore.

BACKGROUND

Two general types of electric propulsion thrusters used in space includeion thrusters and Hall-effect thrusters.

Both ion thrusters and Hall-effect thrusters are ‘electrostatic’electric propulsion devices used for spacecraft propulsion that createthrust by accelerating ions. The thrust created is very small comparedto conventional chemical rockets, but a very high specific impulse, orhigh exhaust velocity, which reduces the propellant requirements formissions is obtained. This high ‘propellant efficiency’ is achievedthrough the very frugal propellant consumption of the electricpropulsion system. They do however require large amounts of power;typically 1 kWe per 0.030-0.040 Newtons thrust for ion thrusters, and 1kWe per 0.050-0.080 Newtons thrust for Hall-effect thrusters.

Ion thrusters and Hall-effect thrusters both generate a beam of ions(electrically charged atoms or molecules) to create thrust in accordancewith Newton's third law. The method of accelerating the ions varies, butall designs take advantage of the charge/mass ratio of the ions. Thisratio means that relatively small potential differences can create veryhigh exhaust velocities. This reduces the amount of reaction mass orfuel required, but increases the amount of specific power requiredcompared to chemical rockets. Electric propulsion thrusters aretherefore able to achieve extremely high specific impulses.

The drawback of the low thrust is low spacecraft acceleration becausethe mass of current electric power units is directly correlated with theamount of power required. This low thrust makes electric propulsionunsuited for launching spacecraft into orbit, but they are ideal forin-space propulsion applications.

Gridded electrostatic ion thrusters commonly utilize xenon gas. This gashas no charge and is ionized by bombarding it with energetic electrons.These electrons can be provided from an electron source as a hot cathodefilament, or more typically a hollow cathode assembly (HCA), which arethen accelerated in the electrical field of the cathode fall to theanode (Kaufman type ion thruster).

The positively charged ions are extracted by an extraction systemconsisting of 2 or 3 multi-aperture grids. After entering the gridsystem via the plasma sheath the ions are accelerated due to thepotential difference between the first and second grid (named screen andaccelerator grid) to the final ion energy of typically 1-2 keV, therebygenerating the thrust. Typical ion velocities are in the range of20,000-50,000 m/s, and higher for some energetic mission applications.

In spacecraft propulsion, a Hall-effect thruster also accelerates ionsby an electric field. Hall-effect thrusters trap electrons in a radialmagnetic field and then use the electrons to ionize propellant,efficiently accelerate the ions to produce thrust, and neutralize theions in the plume.

The essential working principle of the Hall-effect thruster is that ituses an electrostatic potential to accelerate ions up to high speeds butdoes so without the application of a gridded extraction system used inion thrusters. In a Hall-effect thruster the attractive negative chargeis provided by an electron plasma at the open end of the thrusterinstead of a grid. A radial magnetic field of a few tens of milli-Teslais used to confine the electrons, where the combination of the magneticfield and an attraction to the anode upstream surface force a fastcirculating electron current around the axis of the thruster and only aslow axial drift towards the anode occurs.

A propellant, such as xenon gas is fed through the anode, which hasnumerous small holes in it to act as a gas distributor. As the neutralxenon atoms diffuse into the channel of the thruster, they are ionizedby collisions with high energy circulating electrons.

The xenon ions are then accelerated by the electric field between theanode and the cathode. The ions quickly reach speeds of around 15,000m/s for a specific impulse of 1,500 seconds (15 kN·s/kg). Upon exitinghowever, the ions pull an equal number of electrons with them, creatinga plume with no net charge. The axial magnetic field is designed to bestrong enough to substantially deflect the low-mass electrons, but notthe high-mass ions which have a much larger gyroradius and are hardlyimpeded. About 30% of the discharge current is an electron current whichdoes not produce thrust, which limits the energetic efficiency of thethruster; the other 70% of the current is in the ions. The ionizationefficiency of the thruster is thus around 90%.

The magnetic field thus ensures that the discharge power predominatelygoes into accelerating the xenon propellant and not the electrons, andthe thruster turns out to be reasonably efficient.

Because of the counter-flowing electron and ion currents in theHall-effect thruster channel, a greater ion flux can be achieved ascompared to that of the ion thruster—thereby yielding higherthrust-to-power than ion thrusters. Ion thrusters however are capable ofachieving higher exhaust velocities with higher overall thrustefficiencies.

BRIEF SUMMARY

It would be desirable in the propulsion field to provide an electricpropulsion device that delivers the performance capabilities of highthrust-to-power devices (such as Hall-effect thrusters) and the highspecific impulse of high total-impulse devices (such as electrostaticgridded-ion thrusters).

One limitation is that no single Electric Propulsion Thruster (EPT)exists which can operate over the full specific impulse range ofinterest and do so with the combined characteristics of both Hall-effectand ion thrusters, such as high Thrust-to-Power (T/P).

The design attributes of the Hall-effect thruster and the ion thrustersspecific to their ion-acceleration systems: the Hall-effect thrusterutilizes backstreaming electrons accelerated from an external cathodetoward an anode upstream of a radial-geometry magnetic field within anazimuthally-symmetric channel to generate a plasma and createcounter-flowing accelerated ion current; and the ion thruster utilizes aclosely-spaced multi-aperture electrodes (electrostatic ‘ion optics’)with a large applied E-field to focus and accelerate ions from adischarge plasma to form a (space-charge-limited)˜mono-energetic beam.

DESCRIPTION OF DRAWINGS

The accompanying drawings, which are incorporated in and constitute apart of the specification, illustrate various example systems, methods,and so on that illustrate various example embodiments of aspects of theinvention. It will be appreciated that the illustrated elementboundaries (e.g., boxes, groups of boxes, or other shapes) in thefigures represent one example of the boundaries. One of ordinary skillin the art will appreciate that one element may be designed as multipleelements or that multiple elements may be designed as one element. Anelement shown as an internal component of another element may beimplemented as an external component and vice versa. Elements may not bedrawn to scale and in some instances, cross-hatching is not shown toimprove clarity.

FIG. 1 is a perspective view of selected components of an exploded,partially cut-away, electric propulsion machine.

FIG. 2 is a side view of ion optics.

FIG. 3 is a cross-sectional view taken along line 3-3.

FIG. 4 is a side cross-sectional view of an electric propulsion machine.

FIG. 5 is a top plan view, partially cut away, of an exemplary electricpropulsion machine.

FIG. 6 is a side cross-sectional view of the electric propulsion machineshown in FIG. 5.

FIG. 7 is a simplified electric power circuit for an electric propulsionmachine.

FIG. 8 is a side cross-sectional view of an alternate manifestation ofan electric propulsion machine.

FIG. 9 is a simplified electric power circuit for an electric propulsionmachine.

FIG. 10 is a side cross-sectional view of an alternate manifestation ofan electric propulsion machine.

FIG. 11 is a chart illustrating operating modes of an exemplary electricpropulsion machine.

FIG. 12 is a plot illustrating maximum input power and specific impulsefor an exemplary electric propulsion machine.

FIG. 13 is a plot illustrating maximum thrust and specific impulse foran exemplary electric propulsion machine.

FIG. 14 is a plot illustrating input power and specific impulse for anexemplary electric propulsion machine.

FIG. 15 is a plot illustrating thrust and specific impulse for anexemplary electric propulsion machine.

FIG. 16 is a plot illustrating engine efficiency vs. specific impulse ofan exemplary electric propulsion machine.

DETAILED DESCRIPTION

As used herein, the relevant design attributes of a Hall-effect and ionthrusters are their respective ion-acceleration systems: anazimuthally-symmetric channel with axial E-field and radial B-field forthe Hall-effect thruster yielding closed-drift electrons to generate aplasma and create counter-flowing accelerated ion current; andclosely-spaced multi-aperture electrodes or electrostatic ion opticswith a large applied E-field to focus and accelerate ions from adischarge plasma to form a space-charge-limited mono-energetic beam forthe ion thruster.

The plasma production and acceleration mechanisms of the Hall thrusterare closely-coupled and are intimately connected to the geometricconstruction of the thruster discharge geometry.

On the other hand ion thrusters have de-coupled plasma production andacceleration mechanisms. As such, the thruster discharge geometry can beconstructed in a variety of fashions without compromising theoperational integrity of the ion thruster—so long as the accelerationmechanism, the electrode geometry, is maintained. An ion thrusterdischarge chamber may take a number of geometries—cylindrical, anoblate-spheroid, rectangular-box, etc. So long as there is a high degreeof azimuthal symmetry to the ion thruster discharge geometry, a magneticcircuit can be designed to contain the discharge plasma that will yielda high discharge electrical efficiency. This is particularly true whenthe thruster is operated at high plasma densities. Maintaining highdischarge electrical efficiency is an important consideration whentrying to improve thrust to power (T/P)-ratio and overall thrusterefficiency.

In a first embodiment illustrated in exploded and partially cut awayFIG. 1, an electric propulsion machine includes an ion thruster 100comprising an annular discharge chamber 110 and annular ion optics 120covering an exhaust annulus 130. The ion thruster further includes acentrally-mounted neutralizer cathode 140 all arranged along a commonaxis A. Off axis A, but disposed within the annular discharge chamber110, an exemplary ion thruster further includes at least one dischargecathode 150 and a propellant source 160 together forming a dischargecathode assembly 170. As further discussed below, an anode (not shown)is disposed within the annular discharge chamber 110.

It is appreciated that the ion optics 120 are shown “exploded.” from theexhaust annulus 130. The ion optics 120 may be configured as a set ofparallel annular electrodes having an outer radius 180 substantiallyconforming to an outside edge of the ion thruster 100. The ion opticsfurther are defined by an inner radius 182 selected to ensure theexhaust annulus is covered by the ion optics 120. A distance or span 190is defined between the outer radius 180 and the inner radius 182.Moreover, ion optics 120 include closely-spaced apertures 122, usuallycircular, through the thickness which are aligned between theelectrodes. In one example, on the outer most electrode, the aperturesare 0.075″ diameter, with 0.093″ center-to-center spacing in a hexagonalarray, so the electrode has a very high open area fraction. Using theexample, across any 1-inch span there are ˜10 apertures.

With reference now to FIGS. 1 and 2, a side view illustrates that theion optics 120 may include a set of two parallel, substantially planarannular electrodes, 210, 220. The electrodes are spaced apart by asubstantially uniform gap 230. In an embodiment the electrodes 210, 220comprise flat pyrolytic graphite ion optics.

With reference now to FIG. 3, a top-down, cross-sectional view takenalong III-III of FIG. 1, illustrates one embodiment of a plurality ofdischarge cathode assemblies 170 distributed about a closed end 310 ofthe annular discharge chamber 110 opposite of the exhaust annulus.Propellant channel 320 provides a path of fluid communication to theneutralizer cathode 140. In another embodiment, the discharge chambermay be defined by a cylindrical shape 350 on an exterior side and aconic shape (not shown) on the interior producing a chamber of linearlyincreasing or decreasing annular, cross-sectional spans between theclosed end to the exhaust annulus.

With reference now to FIG. 4, cross-sectional, side view of an exemplaryelectric propulsion machine configured as an ion thruster 400 shows anannular discharge chamber 410 and annular ion optics 420 covering anexhaust annulus 430. The ion thruster 400 further includes a centralcylinder 440 defining an interior surface 442 of the annular dischargechamber 410. The interior surface may be configured with magneticshielding 472 to create a field-free region along the central core.Opposite the exhaust annulus 430, the discharge chamber 410 terminatesin a closed annular end 444. An outermost surface 446 opposes theinterior surface 442 and together, interior surface 442, closed end 444and outermost surface 446 define the discharge chamber 410.

An anode 450 may be operatively disposed within and electricallyinsulated from the discharge chamber 410. In the illustrated embodiment,the anode 450 is electrically isolated and disposed along the interiorsurface 442, closed end 444 and outermost surface 446. Such an anode 450provides additional surface area when compared to an anode disposedalong only a single surface, for example the outer surface. As morecompletely described below, the added surface area improves electricalcharacteristics of the engine. Ion thruster 400 is further illustratedwith a discharge cathode assembly 460 to generate a discharge plasma,the cathode assembly including a cathode 462 and propellant feed 464.Rare earth permanent magnets 470 are embedded within or formedintegrally with the interior surface 442 and outermost surface 446 toestablish a boundary magnetic circuit, of ring-cusp, or line-cuspgeometry, to contain and control the energetic electrons emitted fromthe cathode assembly 460 to create an efficient plasma. Although onlyillustrated partially on a single side of the chamber 410, it isunderstood that the magnets 470 line the surfaces 442, 446. At least oneion plenum 480 is disposed toward the exhaust annulus 430 to provide asecond, and primary, propellant feed.

One advantage of the illustrated ion thruster 400, as compared to ionthrusters of conventional configuration (cylindrical discharge withspherically-domed circular ion optics), includes the annular dischargechamber being able to provide for efficient packaging by providing acentral position for mounting the neutralizer cathode assembly (NCA) 486within the annulus. This reduces the outer profile of the engine andeliminates the need for a cantilevered-outboard NCA employed onconventional ion thrusters.

Another advantage of the illustrated ion thruster 400 includes theannular-geometry ion optics allowing for scaling of ion thrusters tovery high power by permitting very-large beam areas with relativelysmall electrode spans, and relatively small span-to-gap ratios. Thisreduces the manufacturing, and the mechanical and thermal stabilityissues inherent with attempting to increase the beam area via increasingthe diameter of spherically-domed ion optics as on conventionalcylindrical ion thrusters.

Another advantage of the illustrated ion thruster 400 includes theannular-geometry ion optics allowing for the application of flat ionoptics electrodes. Flat electrodes improve thrust to power (T/P)-ratiosand efficiencies as compared to conventional ion thrusters bysubstantially reducing or eliminating the off-axis beam vectoring ofions which occurs with spherically-domed ion optics electrodes.

Another advantage of the illustrated ion thruster 400 includes theannular-shaped discharge chamber providing an opportunity to increasethe effective anode-surface area for electron-collection as compared toa conventional cylindrically-shaped ion thrusters of equivalent beamarea. This allows the illustrated engine to operate at thefull-capability of the ion optics, and not have its maximum input powerlevel limited by the available anode surface area. This increase inanode surface area allows the engine to operate at higher dischargecurrents and therefore higher beam currents and input power levels thana conventional ion thruster of equivalent beam area for a given specificimpulse. Alternatively, for the same input power, the increased anodesurface area associated with this annular geometry allows the engine toinclude a smaller outside diameter than a conventional ion thruster.

For example, the NEXT ion thruster (nominal 40 cm beam diameter) has abeam area of approximately 1257 cm², and a discharge chamber anode areaof approximately 3334 cm². An engine according to the teachings herepermit a smaller diameter (comparable to that of the 30 cm diameterNSTAR ion thruster) yielding a comparable discharge chamber anode areaof approximately 3336 cm², using an outside annular discharge chamberdiameter of only 31 cm and an inside annular discharge chamber diameterof 6 cm (inside which a neutralizer cathode assembly may be contained).As detailed in Table 1, an annular discharge chamber engine may supportan annular beam area of about 727 cm² which is sufficient to supportoperation of the engine at the full-power operating power of the largerNEXT thruster—3.52 A beam current and 6.86 kWe. Such an engine wouldhave a much reduced optics span, and optics span-to-gap ratio ascompared to the NEXT ion thruster.

TABLE 1 Engine Conventional EPT Attribute NEXT Ion Thruster AnnularEngine Beam Diameter, cm 40.0 31.0 Beam Area, cm² 1257 727 DischargeChamber 3334 3336 Anode Surface Area, cm² Anode Surface Area to 2.56:14.59:1 Beam Area Optics Span, cm 40.0 12.5 Optics Span-to-Gap Ratio 606:1  189:1

In operation an annular thruster channels propellant flow through thedischarge cathode assembly (DCA) 460, the ion plenum 480, and thecentral common NCA 486. As discussed more completely below, an Ion AnodePower Supply, an Ion Beam Power Supply and an Ion Accelerator ElectrodePower Supply are all energized. Operating range would be typically2000-5000 seconds specific impulse.

With reference now to FIG. 5, in another embodiment, a dual electricpropulsion machine includes an ion thruster 500 comprising an annulardischarge chamber 505 underlying ion optics 510 and a second thruster520, illustrated and described here as a Hall-effect thruster. In thisillustration an ion thruster 500 including an annular-shaped dischargechamber and annular-shaped ion optics 510 are mounted circumferentiallyaround the exterior of a second thruster 520, a Hall-effect thrusterarranged on a common axis and a centrally-positioned neutralizer cathode540 common to both the ion and Hall-effect plasma sources.

Referring now to FIG. 6, a cross-sectional, side view of the exemplarydual electric propulsion machine shown in FIG. 5 includes an annular ionthruster component 600 surrounding a central, second electric propulsionthruster component configured as a Hall-effect thruster 610. The annularion thruster 600 includes an annular discharge chamber 620. The annulardischarge chamber 620 is defined by an outer surface 622, a closedannular end 624 and an interior surface 626 that may be configured withmagnetic shielding to magnetically shield the NCA from the magneticcircuit of the ion thruster and Hall-effect components 600, 610. Theannular ion thruster 600 further includes annular ion optics 630covering an exhaust annulus 640, and an anode 650 operatively disposedwithin and electrically insulated from the discharge chamber 610. Whilethe illustrated anode 650 is configured to maximize surface area byoccupying substantially the surface area of the outer surface 622, theclosed end 624 and the interior surface 626, alternately structuredanodes may occupy less of the interior space of the annular dischargechamber 620 depending on electrical and power requirements.

The annular ion thruster 600 is further illustrated with a dischargecathode assembly 660 to generate a discharge plasma, the cathodeassembly including a cathode 662 and propellant feed 664. Permanentmagnets 670 are embedded within the outer wall 622 and interior wall626. As above, the magnets are only illustrated partially, it isunderstood that the magnets 670 may be annularly disposed within thewalls 622, 626, or discreet magnets may be spaced around the dischargechamber 620 and/or spaced between the exhaust annulus 640 and the end624. Additionally, at least one ion plenum 680 is disposed near theexhaust annulus 640 side of the chamber 620 to provide a secondpropellant feed. In operation, the annular ion thruster 600 providesthrust from the exhaust annulus 630 as indicated by the arrow 690although it is appreciated that thrust is provided from all orsubstantially all of the exhaust annulus 640.

With continued reference to FIG. 6, the Hall-effect thruster component610 may include a separate anode 692 and plenum 694 to supply propellantto the Hall closed drift channel 696 having a radial magnetic field andaxial electric field to generate plasma and create counter flowingaccelerated ion currents. Electron trajectories are illustrated byarrows emanating from a common neutralizer cathode 698 and thrust isillustrated by arrow 699 but understood to be distributed around theHall-effect component 610.

In one embodiment, a dual electric propulsion machine provides forefficient packaging to minimize overall engine diameter.

In one embodiment of a dual electric propulsion machine using aHall-effect thruster component 610 as the second propulsion engine,maintains the design integrity of the Hall-effect thruster to ensure itsperformance characteristics of high T/P-ratio at low specific impulse;

A centrally-located neutralizer cathode assembly (NCA) 698 may providedual functionality as both the discharge plasma generation and beamneutralization for the Hall-effect component 610, and beamneutralization for the ion thruster component 600. Additionally, acentrally located NCA 698 may reduce the outer profile of the engine andeliminate the need for a cantilevered-outboard NCA (not shown) dedicatedfor the operation of the ion thruster.

The annular-geometry ion optics 630 allow for very-large beam areaswhile creating very small electrode spans, and very small span-to-gapratios. This permits use of flat ion optics electrodes. Flat electrodesyield improved T/P-ratios and efficiencies as compared to Current ionthrusters. This is because conventional ion thrusters are cylindrical ingeometry, requiring spherically-domed ion optics electrodes to ensureboth adequate stiffness for launch vibration, and thermo-mechanicalstability under thermal loads during operation to maintain a uniform,controlled inter-electrode gap over very-large spans. The domedelectrodes however result in thrust-losses associated with beamletsdirected off-axis. Reducing thrust-losses associated with off-axis beamvectoring by using flat electrodes result in an improvement in overallefficiency for a given input power as more completely described below.

The annular-shaped discharge chamber 620 of the ion component 600 of thedual electric propulsion machine increases the effective anode-surfacearea for electron-collection as compared to a conventionalcylindrically-shaped ion thruster of equivalent beam area. This allowsthe ion thruster component 600 to operate utilizing the full-perveancecapability of the ion optics 630, and not have its maximum input powerlevel limited by the available anode surface area. This increase inanode surface area allows the ion component to operate at higherdischarge currents and therefore higher beam currents and input powerlevels than current ion thrusters of equivalent beam area for a givenspecific impulse. This increase in input power capability will be morecompletely described below.

The increase in operable surface area due to expanded surface availablein the annulus also creates more radiative surface, permitting operationat higher discharge power levels as compared to a comparable beam areathruster of conventional construction. This is expected to maintainacceptable temperature margins on critical components such as therare-Earth permanent magnets.

Geometric differences in ion optics electrodes 630, and dischargechamber anode surface areas as compared to conventional ion thrustersare documented in Tables 2 and 3 respectively. For purposes ofillustration, the center core, or second thruster of the dual thrusterengine is assumed to be a NASA GRC 300M Hall-Effect thruster in thisembodiment. The 300M is a 20 kW-class laboratory electric propulsionthruster with an external diameter of 15⅜″ (39 cm).

Dual thruster engines with an ion thruster beam area equivalent to aconventional cylindrical ion thruster are listed in Table 2. As notedthe dual thruster engine approach allows for a dramatic reduction inoptics span, and span-to-gap ratio for a given beam area. Specifically,in one example, a 4-6× reduction in span and span-to-gap ratios. Thiscomparatively small span is expected to result in a first-mode naturalfrequency high enough to allow for the use of flat electrodes. This incombination with the application of a long-life, low thermalcoefficient-of-expansion material such as pyrolytic-graphite for theelectrodes will provide an optimal flat-electrode design solution whichis expected to eliminate the thrust-losses inherent with a domedelectrode geometry used in conventional ion thrusters.

TABLE 2 Engine Conventional EPTs Dual Thruster Engine Optics AttributeNSTAR NEXT 50 cm A B C Area, cm² 625 1257 1963 625 1257 1963 Span, cm28.2 40.0 50.0 4.56 8.43 12.2 Span-to-Gap Ratio 427:1 606:1 757:1 69:1128:1 185:1

The differences in anode surface areas between conventional electricpropulsion thrusters (Ion thrusters: NSTAR and NEXT are partial-conic;50 cm is cylindrical) and the annular-portion (ion component) of thedual thruster engine having an equivalent beam area are documented inTable 3 (assuming a core of a 300M Hall-Effect thruster component). Asnoted, the dual thruster engine yields a much larger anode surface areafor an equivalent beam area; e.g. −2.4× increase in area as compared tothe NEXT ion thruster. It should be noted that these calculations arebased on the increase in geometric surface area of the anode. Forring-cusp magnetic circuit plasma discharges the actual effective anodesurface area is the sum of the magnetic cusp lineal areas. Whileestimated, the surface area ratios documented in Table 3 should bereasonably accurate.

TABLE 3 Engine Conventional EPTs Dual thruster Engine Attribute NSTARNEXT 50 cm ‘A’ ‘B’ ‘C’ Beam Area, cm² 625 1257 1963 625 1257 1963Discharge Chamber 2138 3334 5490 6776 7954 9192 Anode Surface Area, cm²Anode Surface Area to 3.42:1 2.65:1 2.80:1 10.84:1 6.33:1 4.68:1 BeamArea Ratio of Dual Thruster  3.17:1 2.39:1 1.67:1 Engine-to-Conventional EPT Anode Area

The comparatively-large surface area of the dual thruster engine anodecompared to either beam area alone or anode area of conventional EPT'sof the same or similar beam areas allows for operation at higherdischarge currents. This enables operation at much higher power levels,thereby taking full-advantage of the current-extraction capability(perveance) of the ion optics.

Several other design features and attributes of embodiments of the dualthruster-engine are noted here, with continued reference to FIG. 6.

The Hall-effect component 610 of the dual thruster engine may be ofconventional construction, using a solenoid electromagnet (not shown) tocreate the appropriately-shaped radial magnetic field at the exit planeof the channel. Use of a centrally-located cathode 698, andimplementation of the extensible-channel concept (to enhance life time)employed on the NASA GRC HiVHAC low power Hall-effect thruster may beused in the dual thruster.

A ring-cusp magnetic circuit is shown within the annular ion dischargechamber 620, created by rows of alternating-polarity rare-Earthpermanent magnets 670 attached or embedded within the surfaces of theannular discharge chamber surfaces 622, 626. Magnetic shielding 676 maybe added between the exterior of the Hall-effect discharge 678 and theion exhaust annulus 640 to separate the magnetic circuits of the twodischarges. Alternately, it may be possible to use the fringe-magneticfield created by the Hall-effect discharge solenoid magnet—which wouldnaturally penetrate into the annular ion discharge chamber 620 withoutshielding—and shape it by appropriate application of magnetic materialswithin the walls of the annular ion discharge chamber 620 to generateand control the discharge plasma in this zone. This would be with thepotential benefit of reduced mass by elimination of the rare-Earthpermanent magnets 670.

The annular discharge chamber 620 of the ion thruster component 600 mayinclude a conventional discharge cathode assembly (DCA) 662 to generatethe discharge plasma.

Typical ion thrusters require 3 separate propellant feeds; a Hall-effectthruster 2 propellant feeds. In one embodiment a dual thruster uses atotal of 4 separate propellant feeds: one each for the ion discharge 680and Hall discharge 694; one for the ion DCA 664; and one for the centralcommon NCA 698.

The annular discharge chamber 620 of the ion thruster component 600 mayuse a ‘reverse-feed’ plenum which may inject the propellant from the ionoptics-end of the discharge backwards, giving it an initial axialvelocity component resulting in an increased neutral atom residence timeand improved propellant efficiency.

The dimensions of the channel width and depth of the Hall-effectcomponent 610 would be defined by the intended operating power level(s).

The annular area of the ion optics 630 would be established by theintended operating power level(s), which would then establish theelectrode span (and discharge channel width) and overall dual thrusteroutside diameter.

The depth of the annular ion component discharge chamber 620 may includea tradeoff of electrical and propellant efficiencies, maximum desiredinput power, and overall dual thruster mass. It may be, for example,most effective mechanically and magnetically to match the channel depthsof the ion thruster component 600 and the Hall-effect component 610.

With reference now to FIG. 7, a simplified electrical diagram for a dualthruster is shown. Of note is that sufficient commonality exists in therequirements associated with the ion thruster component 600 beam powersupply and the Hall-effect component 610 anode power supply that thisfunction could be performed by a common power converter.

With reference now to FIG. 8, in another embodiment of the dualthruster, it may be advantageous to eliminate the DCA, and rely on thecentral common NCA 810 for plasma generation and beam neutralization ofboth the Hall-effect thruster component 820 and ion thruster component830 of the dual thruster. Such a device may comprise one hollow cathodeemitter 812 and propellant feed 814, reducing the total propellant feedsfor the engine from 4 to 3.

One additional advantage to this approach is to eliminate the asymmetryin the ion thruster component 830 created by a singular dischargecathode assembly within an annular-shaped discharge chamber 836. Thisasymmetry could cause a (minor) asymmetry in the plasma at the plane ofthe ion optics 840, and also result in a (minor) asymmetry in themagnetic circuit (‘a hole’) in the area of the DCA resulting in anincrease in discharge electrical losses.

In the same manner that the discharge plasma is generated for theHall-effect component, namely electron-back-streaming of current fromthe NCA 810 to the Hall-effect anode 846, it is expected that adischarge plasma for the ion thruster component 830 may be generatedsimilarly. In the embodiment illustrated in FIG. 8, radial slots 850 areprovided around the circumference of the exterior of the Hall-effectcomponent 820 to provide a passage to the annular-shaped dischargechamber 836 for neutrals, ions, and energetic electrons.

By appropriate shaping of the magnetic field in the passage (e.g. —axialB-field component across the radial slots) and de-energizing theHall-effect solenoid (not shown) creating the downstream radial B-fieldcomponent in the Hall channel, it should be possible to back-streamelectrons 860 from the NCA 810, through the radial slots 850 toward theion discharge anode 870 and in the process deplete their energy and usethem efficiently to generate the ion thruster discharge plasma. Thelocation of electron current collection and hence which discharge isoperated (Hall-effect or ion) could be controlled by the switchesidentified in FIG. 7 and by the propellant flow rates through therespective plena. Alterations are available to artisans to optimize themagnetic field of the ion discharge may be accomplished to meet designconstraints.

With reference now to FIG. 9, a simplified electrical schematic for adual thruster embodiment having a single, central neutralizer cathodeassembly 810 is shown. Although there is no longer a separate DCAdischarge, the ion anode power supply 920 remains to provide a biasvoltage to the screen electrode 930 of the ion optics 840.

With reference now to FIG. 10, in yet another embodiment, a dual enginethruster 1000 includes an ion thruster component 1010 and a Hall-effectcomponent 1020. While similar in many respects to the embodiments above,one distinction lies in the ion discharge chamber 1030 and an attempt toeliminate asymmetries caused by off-axis discharge cathode assemblies inthe annular discharge chamber. Specifically, discharge chamber 1030includes an annular section 1032 and a cylindrical section 1034. Annularsection 1032 is bordered on one side by the exhaust annulus 1040 and bythe cylindrical section 1034 on the opposing side. Desirably, a singledischarge cathode assembly 1050 including cathode 1052 and propellantfeed 1054 lies on central axis A. To support the Hall-effect component1020, a mechanical support 1060 is provided, although it is appreciatedthat other supports in addition to the illustrated mechanical supportcould be interchanged. Supporting floor 1062 of mechanical support 1060is insulated from the annular ion anode (discussed below) and is held atcathode potential.

The illustrated embodiment includes an annular ion anode 1070 ofpotentially greater surface area permitting higher currents anddesirable features as described above. In the embodiment shown, annularion anode 1070 lines but is insulated from an exterior surface 1082, abottom 1084 of the cylindrical section, and an interior surface 1086.

As described in the embodiments above, a dual thruster is anticipated tohave certain operational and performance characteristics. For example, adual thruster can be operated in multiple ‘modes.’ Referring now to FIG.11, in a ‘Hall-effect mode’ 1110 the propellant flow is channeledthrough the Hall plenum and the central common NCA. The solenoid togenerate the radial magnetic field component and the IonBeam/Hall-Effect Anode Dual Power Supply are energized. S1 switch isopen, and S2 switch is closed. The Hall-effect component of the dualthruster is then operated as a conventional Hall thruster. Operatingrange would be typically 1200-2000 seconds specific impulse as seen at1110.

As another example, the dual thruster can be operated in an ‘ion mode’1120 where the propellant flow is channeled through the ion plenum, theDCA, and the central common NCA. The Ion Anode Power Supply, the IonBeam/Hall-Effect Anode Dual Power Supply, and the Ion AcceleratorElectrode Power Supply are all energized. S1 switch is closed, and S2switch is open. The ion component of the dual thruster is then operatedas a conventional ion thruster. Operating range would be typically2000-4000 seconds specific impulse as seen at 1120.

In yet another example, the dual thruster can be operated in a ‘burstmode’ 1130 where the propellant flow is channeled through all 4locations, all power supplies are energized, and both S1 and S2 switchare closed. The Hall-effect and ion components of the dual thruster arethen both operated simultaneously. Operation would be typically in the1800-2200 seconds specific impulse range, in a zone where bothcomponents are capable of functioning with some overlap in capability1130. This mode is theoretically possible but may not be operationallyadvantageous. However, two potential reasons for operating in this burstmode include: (a) Providing a seamless-transition in specific impulsethrottling between Hall-effect mode 1110 and ion mode 1120 operation;and (b) assuming there is sufficient capability in the power electronicsand propellant management system, operating at a total input power andgenerating a total thrust level for the dual thruster exceeding thatwhich could be achieved by either the Hall-effect or ion componentsalone.

As mentioned above in discussing the annular discharge chamber ionthruster, when the dual thruster is operated in ion mode 1120 higherefficiencies at fixed input power, and higher input power at fixedspecific impulse are possible, as compared to conventional ionthrusters. For example, the flat-geometry of the ion optics electrodesafforded by the annular design will improve the efficiency of thethruster as compared to a conventional ion thruster of equivalent beamarea. This is because the thrust-losses associated with beam divergencedue to the domed shape of conventional ion thrusters are eliminated inthe annular design.

Moreover, both the specific impulse and the thrust are proportional tothe thrust-loss correction factor due to off-axis beam vectoring(F_(t)); hence the overall thruster efficiency, which includes thespecific impulse and thrust terms, is proportional to F_(t) ². Thecorrection F_(t) includes both the beam divergence due to the electrodedome shape F_(t-d) and beam divergence due to beamlet expansion F_(t-b).

From equation 9 of Soulas, G. C., “Design and Performance of 40 cm IonOptics,” the NEXT ion thruster optics (dome height of approximately 2.35cm and chord of 18 cm) F_(t-d) is estimated to be approximately 0.983.Hence, the reduction in thrust and specific impulse, and overallthruster efficiency of the NEXT ion thruster due to off-axis beamvectoring caused by the domed ion optics is expected to be 0.017, or−1.7% {100*(F_(t-b)−1.00)} in thrust and specific impulse, and 0.034, or−3.4% {100*(F_(t-b) ²−1.00)} in efficiency at all throttle conditions.Therefore, all else being equal, for an equivalent beam area dualthruster with flat ion optics (F_(t-d) equal to 1.00), a 3.4% increasein efficiency as compared to the NEXT ion thruster would be expectedacross the entire specific impulse range as seen below:

TABLE 4 Throttle Thruster level/Input NEXT-STEP & NEXT Annular IonThruster Power T/P- Specific T/P- Specific Level, ratio, Impulse, ratio,Impulse, Effi- kW mN/kW sec Efficiency mN/kW sec ciency —/2.498 60.81711 0.499 61.8 1740 0.516 —/3.658 54.9 2240 0.601 55.8 2277 0.621—/4.818 50.0 2666 0.655 50.8 2711 0.677 TL23/2.816 42.2 3090 0.640 42.93142 0.662 TL25/3.683 37.8 3616 0.670 38.4 3676 0.693 TL12/2.439 32.93999 0.645 33.4 4066 0.667 TL40/6.860 34.5 4188 0.708 35.1 4258 0.732

One of the design issues with the NEXT thruster, and other conventionalpartial-conic or cylindrical discharge chamber ion thrusters (such asthe NSTAR ion thruster, or the NASA GRC 50 cm laboratory model ionthruster), is that they cannot take full advantage of the ion currentextraction capability of the ion optics technology—and hence operateat-or-near their maximum theoretical input power capability.

The NEXT neutralizer cathode, discharge cathode, magnets, ion optics,and high voltage propellant isolators all have adequate thermal and/oroperational margins for operation at extremely-high input power levels;well in excess of the 7 kW maximum input power of the NEXT throttletable. Additionally, conventional NEXT ion optics are capable ofoperating at beam currents well-in-excess-of the maximum 3.52 A of theNEXT throttle table; >>7.0 A beam currents at full total voltage (2010V). Application of advanced high-perveance design ion optics to the NEXTthruster would allow for operation at beam currents >>7.0 A at low totaland beam voltages, thus enabling truly-high Thrust-to-Power operation.

Unfortunately, although operation at these high beam currents isconsistent with the ion extraction system electrostatic functionality,they require a maximum sustainable discharge current which exceeds thatwhich can be supported by the available anode surface area of the NEXTthruster discharge chamber. The maximum sustainable discharge currentfor the NEXT thruster is estimated to be in the range of about 32-35Amperes, yielding a maximum beam current of about 7.0 A.

As documented in Table 3, for a given beam area, implementation of theannular discharge chamber either in a dual thruster arrangement or as astand-alone annular ion thruster increases the effective and availableanode surface area, as compared to conventional ion thrusters. Thelarger anode surface area increases the permissible sustainabledischarge current and therefore beam current and thruster input powerfor a given beam voltage (specific impulse). The increase in sustainabledischarge current (and approximate equivalent increase in beam currentand input power) will be either directly proportional to the increase inanode surface area—or—will be equal to the discharge current necessaryto support the maximum perveance-limited beam current that the ionoptics are capable of extracting, whichever is less.

For an ion thruster with beam area equivalent to that of the NEXT ionthruster, an increase in anode surface area of 2.4× is possible. Thispermits an increase in sustainable discharge current and increase inbeam current and thruster input power for a given beam voltage (specificimpulse).

The increase in input power for a dual thruster versus a NEXT ionthruster of equivalent beam area is documented in Table 5 for a range ofspecific impulse. As indicated in Table 5 the input power to the NEXTion thruster is limited by the anode-area down to about 3340 secondsspecific impulse. Below this level, both the anode-area and the ionoptics current-extraction-capability limit the thruster input power.

The annular ion and dual thruster is also anode-area limited at highspecific impulse (4430 seconds); the ion optics are capable ofsupporting >20 A beam current, which would require a discharge currentof greater than 84 Amperes. Although at this specific impulse the engineis anode-area limited, the maximum input power and thrust for thisengine are 2.4× higher than that feasible with the equivalent-beam areaNEXT thruster, due to the larger anode area of the dual thruster. At4140 seconds specific impulse and below the dual thruster maximum inputpower is limited by the ion optics current-extraction-capability. Thatis, the engine is much more capable of taking full-advantage of the ionoptics electrostatics than the NEXT thruster. At about 3340 secondsspecific impulse and below, the maximum input power and thrust areequivalent to the NEXT thruster.

TABLE 5 Attribute NEXT Ion Thruster Dual Engine B (Tables 2 and 3) BeamArea, cm² 1257 1257 Discharge 3334 7954 Chamber Anode Surface Area, cm²Maximum ≦35 ≦84 Sustainable Discharge Current, A Maximum Anode- ≦7.0≦16.8 Area-Limited Beam Current, A Anode Max. Max. Max. Area- Max. BeamAnode Area- Input Max. Beam or- Input Max. Specific Impulse Current,or-Optics Power, Thrust, Current, Optics Power, Thrust, sec A Limited kWmN A Limited kW mN 4434 7.04 Anode 13.646 472.1 16.8 Anode 32.501 11274139 7.04 Anode 12.042 440.8 14.5 Optics 24.755 907.8 3908 7.04 Anode10.865 416.2 11.0 Optics 16.951 650.2 3592 7.04 Anode  9.371 382.5  7.30Optics  9.715 396.6 3338 7.04 Anode/Optics  8.287 355.5  7.04 Optics 8.287 355.5 3188 7.04 Anode/Optics  7.689 339.5  7.04 Optics  7.689339.5 3031 7.04 Anode/Optics  7.082 322.8  7.04 Optics  7.082 322.8 27337.04 Anode/Optics  6.027 291.0  7.04 Optics  6.027 291.0

The data in Table 5 for the dual thruster are conservative estimates.This is because: a) the 1.7% increases in both specific impulse andthrust expected with the dual thruster due to the flat-geometryelectrodes were not included; and b) the design of the dual thrusterassumed the same ion optics electrostatic design as that of the NEXTthruster. The dual thruster could in fact incorporate anadvanced-perveance ion optics design in an annular configuration—whilemaintaining the same beam area as the NEXT thruster optics. This dualthruster design would yield an increase in input power and thrust ascompared to that documented in Table 5. It should be noted that therewould be no advantage to applying advanced-perveance ion opticsconfiguration to the NEXT thruster since its maximum beam current isalready limited or inhibited by the anode-area.

With reference now to FIG. 12, a graph illustrates the differences inmaximum input power as a function of specific impulse for the DualEngine B 1210 in Table 5 and an exemplary current thruster 1220 (NEXT)of equivalent beam area and electrode geometry and as set forth in Table5. As can be seen, the slope of the curve 1210 dramatically increasesabove 3500 seconds specific impulse.

With reference now to FIG. 13, a graph illustrates the differences inmaximum thrust capability as a function of specific impulse for the DualEngine B 1310 in Table 5 and an exemplary current thruster 1320 (NEXT)of equivalent beam area and electrode geometry and as set forth in Table5. As can be seen, the slope of the curve 1310 dramatically increasesabove 3500 seconds specific impulse.

It can now be appreciated that many variations of the dual thrusterdescribed here are possible. For example: using a second, or interiorthruster other than a Hall-effect component, combining various sizes ofHall-effect components and ion components, to yield difference totalspecific impulse throttling ranges, and different total ranges in inputpower and thrust capability—depending upon the application need. Onecase example is provided here to define overall performance of a dualthruster. This example is for illustration, with the performancecharacteristics quoted specific to the selected configuration. As suchthe performance numbers documented do not imply any particular limits incapability to the dual thruster concept in general.

Table 6 lists the projected performance capabilities of one embodimentof a dual thruster, with a Hall-effect component having physicalcharacteristics similar to that of the NASA GRC 300M Hall thruster, andan Ion component with beam area equivalent to the NASA NEXT Ionthruster. The performance of the Hall-effect component with specificimpulse was modeled assuming a similar efficiency-specific impulsecharacteristic as the BPT-4000 Hall-Effect thruster, with a nominalinput power of 20 kW at 2000 seconds specific impulse which is thedesign basis for the 300M.

The ion component of the dual thruster was modeled in a fashionconsistent with prior ion thruster performance modeling conducted bythis author assuming a discharge chamber, propellant efficiency of 0.92.The performance gains due to the elimination of the thrust-lossesassociated with domed ion optics were included in these calculations.Additionally, an advanced-perveance ion optics electrode design wasassumed using equation 15 of Patterson, M. J., “NEXT Study of ThrusterExtended-Performance II (NEXT STEP II).

TABLE 6 Thrust- to- Specific Power Impulse, Input Power, Ratio, Mode seckW Thrust, mN mN/kW Efficiency Hall- 1220 4.440 351.1 79.0 0.473 EffectHall- 1440 5.560 408.9 73.6 0.520 Effect Hall- 1610 6.670 457.8 68.70.542 Effect Hall- 1740 7.780 510.0 65.6 0.560 Effect Hall- 1840 13.330844.4 63.3 0.572 Effect Burst 1840 19.201 1196 62.3 0.562 (13.330 +5.871) (844.4 + 351.5) Burst 1960 23.362 1610 61.1 0.587 (20.000 +6.362) (1236 + 374.5) Burst 2150 27.206 1540 56.6 0.597 (20.000 + 7.206)(1129 + 411.0) Ion 2150 7.206 411.0 57.0 0.602 Ion 2207 7.471 421.7 56.40.611 Ion 2298 7.912 439.0 55.5 0.625 Ion 2538 9.172 484.9 52.9 0.658Ion 2758 10.432 526.9 50.5 0.683 Ion 2774 12.188 614.1 50.4 0.685 Ion3076 13.703 648.5 47.3 0.714 Ion 3235 15.634 716.3 45.8 0.727 Ion 338716.875 750.0 44.4 0.738 Ion 3645 19.561 823.6 42.1 0.752 Ion 3966 25.6911010 39.3 0.765 Ion 4200 28.573 1070 37.5 0.771 Ion 4499 32.501 114635.3 0.778

Note in Table 6, ‘Hall-Effect Mode’ refers to the Hall-effect componentoperating solely (from 1220-1840 seconds specific impulse), ‘Ion Mode’refers to the ion component operating solely (from 2150-4500 secondsspecific impulse), and ‘Burst Mode’ refers to both components operatingsimultaneously. In this example case, burst mode is operated in thespecific impulse ‘overlap-zone’ of 1840-2150 seconds specific impulse.The individual input power and thrust contributions of the twocomponents in burst mode are documented in Table 6, with the firstnumber associated with the Hall-effect component and the second numberassociated with the ion component.

As noted in Table 6, this dual thruster provides a continuous-throttlingcapability from 1220-4500 seconds specific impulse (3.7:1 range), withan input power range of 4.44-32.5 kW (7.3:1), thrust of 351-1146 mN(3.3:1), and efficiency of 0.47-0.78. A peak in input power and thrustoccur over the specific impulse overlap-zone of the burst mode, withslight reduction in efficiency in this region.

With reference now to FIG. 14, input power vs. specific impulse for thedual thruster of Table 6 is graphed at 1410. For comparison, input powervs. specific impulse for the maximum input power of a NEXT ion thrusteris graphed at 1420.

With reference now to FIG. 15, thrust vs. specific impulse for the dualthruster of Table 6 is graphed at 1510. For comparison, thrust vs.specific impulse for the maximum input power of a NEXT ion thruster isgraphed at 1520.

With reference now to FIG. 16, engine efficiency vs. specific impulse ofthe dual thruster of Table 6 is graphed at 1610. For comparison, alsoplotted is efficiency data for a current Hall-effect thruster 1620(BPT-4000) and a current ion thruster 1630 (NEXT). It is noted that thedual thruster curve 1610 represents the performance capability of asingle device over the useful impulse range.

As seen in FIG. 16 a significant efficiency gain is expected for thedual thruster during operation in ‘Ion Mode’ (>2150 seconds specificimpulse) as compared to current EPTs. These efficiency gains are due tothe elimination of divergence-losses associated with the ion opticsgeometry, along with other efficiency gains as a consequence of the factthat the dual thruster is operating at much higher input power levels ata given specific impulse.

The dual thruster could be applied to any application for which the highthrust-to-power characteristics of a Hall-effect thruster and highspecific impulse capability of an ion thruster would be advantageous.For example: those Earth-orbital applications requiring both primary andauxiliary electric propulsion functions; Earth-orbital applicationsrequiring both rapid orbit changes and fuel-efficient less-time-criticalorbital changes; and Planetary mission applications requiring bothfuel-efficient high-delta-V transfers and on-orbit high thrust-to-poweroperations.

In one embodiment of a dual thruster, the Hall-effect or other suitablesecond thruster component and the ion component may have approximatelythe same input power capabilities at their nominal design specificimpulses (about 2000 seconds for Hall-effect and 3600 seconds for ion).This geometry is referred to here as a ‘Matched-Dual-Mode.’ Thisconfiguration is that provided in the example case earlier, and would beappropriate if both components were required to perform aprimary-propulsive application.

In an alternative embodiment, referred to here as a ‘Mixed-Dual-Mode,’involves combining a Hall-effect or other suitable second thrustercomponent and an ion component with markedly-different input powercapabilities at their respective nominal design specific impulses. Thismay be most-appropriate if one component were required to provideprimary propulsion and the other component provide auxiliary propulsion;for example, combining a high-power Hall-effect component core (such asthe NASA GRC 300M) for rapid (Earth) orbital transfers with a low-powerion component (for example, beam area equivalent to the NASA NSTARthruster) for station-keeping. These configuration options are capturedin Table 7.

TABLE 7 Ion Component Intermediate- Low-Power Power High-PowerHall-Effect Low-Power Matched Mixed Mixed Component Intermediate- MixedMatched Mixed Power High-Power Mixed Mixed Matched

While the systems, methods, and so on have been illustrated bydescribing examples, and while the examples have been described inconsiderable detail, it is not the intention of the applicants torestrict or in any way limit the scope of the appended claims to suchdetail. It is, of course, not possible to describe every conceivablecombination of components or methodologies for purposes of describingthe systems, methods, and so on provided herein. Additional advantagesand modifications will readily appear to those skilled in the art.Therefore, the invention, in its broader aspects, is not limited to thespecific details, the representative apparatus, and illustrativeexamples shown and described. Accordingly, departures may be made fromsuch details without departing from the spirit or scope of theapplicants' general inventive concept. Thus, this application isintended to embrace alterations, modifications, and variations that fallwithin the scope of the appended claims. Furthermore, the precedingdescription is not meant to limit the scope of the invention. Rather,the scope of the invention is to be determined by the appended claimsand their equivalents.

As used herein, “connection” or “connected” means both directly, thatis, without other intervening elements or components, and indirectly,that is, with another component or components arranged between the itemsidentified or described as being connected. To the extent that the term“includes” or “including” is employed in the detailed description or theclaims, it is intended to be inclusive in a manner similar to the term“comprising” as that term is interpreted when employed as a transitionalword in a claim. Furthermore, to the extent that the term “or” isemployed in the claims (e.g., A or B) it is intended to mean “A or B orboth”. When the applicants intend to indicate “only A or B but not both”then the term “only A or B but not both” will be employed. Similarly,when the applicants intend to indicate “one and only one” of A, B, or C,the applicants will employ the phrase “one and only one”. Thus, use ofthe term “or” herein is the inclusive, and not the exclusive use. See,Bryan A. Gamer, A Dictionary of Modern Legal Usage 624 (2d. Ed. 1995).

The invention claimed is:
 1. An electric propulsion thruster comprising:an ion thruster comprising: an annular discharge chamber surrounding aphysically spaced central space, both the annular discharge chamber andthe central space centered around a common axis, the annular dischargechamber having a closed end and an opposite exhaust annulus, thephysically spaced central space extending from the closed end to theopposite exhaust annulus to form an outlet; a discharge cathodeoperatively disposed within the annular discharge chamber; an annularanode operatively disposed within the annular discharge chamber betweenthe closed end and the exhaust end; and ion optics operativelyassociated with the exhaust annulus.
 2. The electric propulsion thrusteras set forth in claim 1, further comprising a neutralizer cathodedisposed along the common axis.
 3. The electric propulsion thruster asset forth in claim 2, further comprising a Hall-effect thruster disposedwithin the central space surrounding the neutralizer cathode.
 4. Theelectric propulsion thruster as set forth in claim 1, where the ionoptics comprise a first substantially planar, annular electrode and aspaced second substantially planar, annular electrode.
 5. The electricpropulsion thruster as set forth in claim 4, where the ion opticscomprise a ratio of a span to a gap less than 300:1, where the spancomprises a distance from an inside radius to a larger outside radius ofthe ion optics and the gap comprises a distance from the first electrodeto the second electrode.
 6. The electric propulsion thruster as setforth in claim 1, where the anode comprises a surface area greater than6000 cm².
 7. The electric propulsion thruster as set forth in claim 1,where a ratio of an anode surface area to beam area is greater than 4:1.8. The electric propulsion thruster as set forth in claim 3, furthercomprising a propellant feed for the ion thruster, a propellant feed forthe Hall-effect thruster, and a propellant feed for the neutralizercathode.
 9. The electric propulsion thruster as set forth in claim 3,further comprising a plurality of radial passages between theHall-effect thruster and the annular discharge chamber.
 10. An electricpropulsion thruster comprising: a neutralizer cathode disposed on acentral axis; a Hall-effect thruster centered on the central axis; andan ion thruster surrounding the Hall effect thruster comprising: adischarge chamber having an exhaust annulus at one end; an anodeoperatively disposed within the discharge chamber; and first and secondsubstantially planar annular electrodes operatively disposed adjacent tothe exhaust annulus, the first and second substantially planar annularelectrodes spaced from each other by a substantially uniform gap. 11.The electric propulsion thruster as set forth in claim 10, where thefirst and second substantially planar annular electrodes comprise aratio of a span to the substantially uniform gap less than 300:1, wherethe span comprises a distance from an inside edge of one of the first orsecond annular electrodes to an outside edge of the same first or secondannular electrodes.
 12. The electric propulsion thruster as set forth inclaim 10, where the anode comprises a surface area greater than 3.5times than that of an area corresponding to a beam area.
 13. Theelectric propulsion thruster as set forth in claim 10, furthercomprising a propellant feed for the ion thruster, a propellant feed forthe Hall-effect thruster, and a propellant feed for the neutralizercathode.
 14. The electric propulsion thruster as set forth in claim 10,further comprising a plurality of radial passages between theHall-effect thruster and the discharge chamber.
 15. The electricpropulsion thruster as set forth in claim 10, where the dischargechamber comprises an annular chamber surrounding the Hall effectthruster in communication with a substantially cylindrical dischargechamber underlying and mechanically supporting the Hall effect thruster.16. The electric propulsion thruster as set forth in claim 10,comprising an engine efficiency greater than 0.70 at a specific impulsegreater than 3000 seconds.
 17. The electric propulsion thruster as setforth in claim 10, comprising an engine efficiency greater than 0.50 ata specific impulse greater than 1500 seconds and an engine efficiencygreater than 0.65 at a specific impulse greater than 3000 seconds. 18.An electric propulsion machine comprising: an ion thruster comprising:an annular discharge chamber surrounding a central cylinder and centeredupon an axis, the annular discharge chamber having an exhaust annulus,the central cylinder extending from a closed end to the exhaust annulusto form an outlet; an annular anode operatively disposed within theannular discharge chamber between the closed end and the exhaust end,the anode including a surface area greater than 3.5 times than that ofan area corresponding to a beam area; and ion optics operativelyassociated with the exhaust annulus.
 19. The electric propulsion machineas set forth in claim 18, further comprising an engine efficiencygreater than 0.70 at a specific impulse greater than 3000 seconds. 20.The electric propulsion machine as set forth in claim 19, furthercomprising an engine efficiency greater than 0.50 at a specific impulsegreater than 1500 seconds and an engine efficiency greater than 0.65 ata specific impulse greater than 3000 seconds.